High power epicyclic gearbox and operation thereof

Abstract

An engine for an aircraft includes an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan having a plurality of fan blades; and a gearbox. The gearbox is an epicyclic gearbox and comprises a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted. The radial bending stiffness of the planet carrier is equal to or greater than 1.20×109 N/m, and/or the tilt stiffness of the planet carrier is greater than or equal to 6.00×108 Nm/rad. A method of operation of such an engine is also disclosed.

The invention claimed is: 
     
         1 . A gas turbine engine for an aircraft comprising:
 an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;   a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and   a gearbox arranged to receive an input from the core shaft and to output drive to the fan via a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, wherein:   a fan diameter is in the range of 360 cm to 420 cm;   a ratio of a radius of a fan blade at its hub to a radius of the fan blade at its tip is less than 0.25;   each fan blade is manufactured at least in part from carbon fibre; and   a torsional stiffness of the fan shaft is equal to or greater than 1.3×10 7  Nm/rad or in a range from 1.3×10 7  Nm/rad to 2.5×10 9  Nm/rad.   
     
     
         2 . The gas turbine engine according to  claim 1 , wherein a mounting structure for the fan shaft comprises a fan shaft support structure, the fan shaft support structure comprising at least two bearings spaced apart along an axial length of the fan shaft, at least one of the bearings being positioned forward of the gearbox and at least one of the bearings being located at a position rearward of the gearbox. 
     
     
         3 . The gas turbine engine according to  claim 1 , further comprising:
 a gearbox support arranged to at least partially support the gearbox in a fixed position within the engine, wherein:   a torsional stiffness of the gearbox support is in a range from 4.2×10 7  Nm/rad to 1.0×10 10  Nm/rad.   
     
     
         4 . The gas turbine engine according to  claim 1 , wherein a torque supplied by the turbine to the core shaft at cruise conditions is in the range from 10,000 to 15,000 Nm. 
     
     
         5 . The gas turbine engine according to  claim 1 , wherein:
 a gear ratio of the gearbox is at least 4;   an overall pressure ratio, defined as the ratio of a stagnation pressure upstream of the fan to a stagnation pressure at an exit of a highest pressure compressor, is in a range from 40 to 55 at cruise conditions;   a bypass ratio is greater than 20 at cruise conditions;   the gas turbine engine is a turbofan engine or an open rotor engine; and/or   a maximum thrust of the engine, measured at an ambient pressure of 101.3 kPa and a temperature of 30 degrees C. with the engine static, is at least 160 kN.   
     
     
         6 . A gas turbine engine for an aircraft comprising:
 an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;   a fan located upstream of the engine core, the fan comprising a plurality of fan blades;   a gearbox arranged to receive an input from the core shaft and to output drive to the fan via a fan shaft so as to drive the fan at a lower rotational speed than the core shaft; and   a gearbox support arranged to at least partially support the gearbox in a fixed position within the engine, wherein:   a fan diameter is in the range of 340 cm to 420 cm;   a ratio of a radius of a fan blade at its hub to a radius of the fan blade at its tip is less than 0.25; and   a torsional stiffness of the gearbox support is equal to or greater than 4.2×10 7  Nm/rad.   
     
     
         7 . The gas turbine engine according to  claim 6 , wherein:
 a torque supplied by the turbine to the core shaft at cruise conditions is in the range from 10,000 to 15,000 Nm; and/or   a maximum thrust of the engine, measured at an ambient pressure of 101.3 kPa and a temperature of 30 degrees C. with the engine static, is at least 160 kN.   
     
     
         8 . The gas turbine engine according to  claim 6 , wherein:
 a gear ratio of the gearbox is at least 4;   an overall pressure ratio, defined as the ratio of a stagnation pressure upstream of the fan to a stagnation pressure at an exit of a highest pressure compressor, is in a range from 40 to 55 at cruise conditions;   a bypass ratio is greater than 20 at cruise conditions;   the gas turbine engine is a turbofan engine or an open rotor engine; and   each fan blade is manufactured at least in part from carbon fibre.   
     
     
         9 . The gas turbine according to  claim 6 , wherein:
 the gas turbine engine is an open rotor engine; and   a mounting structure for the fan shaft comprises a fan shaft support structure, the fan shaft support structure comprising at least two bearings spaced apart along an axial length of the fan shaft, at least one of the bearings being positioned forward of the gearbox.   
     
     
         10 . The gas turbine engine according to  claim 9 , wherein at least one of the bearings is located at a position rearward of the gearbox. 
     
     
         11 . A gas turbine engine for an aircraft comprising:
 an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;   a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and   a gearbox arranged to receive an input from the core shaft and to output drive to the fan via a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, wherein:   a fan diameter is greater than 340 cm;   a ratio of a radius of a fan blade at its hub to a radius of the fan blade at its tip is less than 0.25;   each fan blade is manufactured at least in part from carbon fibre; and   a gear ratio of the gearbox is at least 3.7.   
     
     
         12 . The gas turbine engine according to  claim 11 , wherein a maximum thrust of the engine, measured at standard at an ambient pressure of 101.3 kPa and a temperature of 30 degrees C. with the engine static, is in the range 160 kN to 190 kN, or at least 160 kN. 
     
     
         13 . The gas turbine engine according to  claim 11 , wherein a torque supplied by the turbine to the core shaft at cruise conditions is in the range from 10,000 to 15,000 Nm. 
     
     
         14 . The gas turbine engine according to  claim 11 , wherein:
 an overall pressure ratio, defined as the ratio of a stagnation pressure upstream of the fan to a stagnation pressure at an exit of a highest pressure compressor, is in a range from 40 to 55 at cruise conditions; and/or   a specific thrust, defined as a net thrust of the engine divided by a total mass flow through the engine, is less than 80 NKg− 1 s.   
     
     
         15 . The gas turbine engine according to  claim 11 , wherein:
 the gear ratio of the gearbox is in a range from 4.2 to 4.5; and/or   the fan comprises no more than 14 fan blades.   
     
     
         16 . The gas turbine engine according to  claim 11 , wherein:
 the fan diameter is in the range of 360 cm to 420 cm; and/or   the gear ratio of the gearbox is at least 3.9.   
     
     
         17 . The gas turbine according to  claim 11 , wherein:
 the gas turbine engine is an open rotor engine; and   a mounting structure for the fan shaft comprises a fan shaft support structure, the fan shaft support structure comprising at least two bearings spaced apart along an axial length of the fan shaft, at least one of the bearings being positioned forward of the gearbox.   
     
     
         18 . The gas turbine engine according to  claim 17 , wherein at least one of the bearings is located at a position rearward of the gearbox. 
     
     
         19 . The gas turbine engine according to  claim 11 , wherein:
 an overall pressure ratio, defined as the ratio of a stagnation pressure upstream of the fan to a stagnation pressure at an exit of a highest pressure compressor, is in a range from 40 to 55 at cruise conditions;   the fan diameter is in the range of 350 cm to 400 cm; and   a bypass ratio is greater than 20 at cruise conditions.   
     
     
         20 . The gas turbine engine according to  claim 11 , wherein a ratio of the fan diameter to the number of fan blades is greater than 26.25m, or greater than 30m.